Gas turbine engines are typically used in power plant applications for the purpose of generating electricity. A typical gas turbine engine is comprised of a plurality of combustors, which are arranged in an annular array around a centerline of the engine. The combustors are then provided pressurized air from a compressor of the gas turbine engine. The pressurized air is mixed with fuel and the mixture is ignited to produce high temperature combustion gases. These high temperature combustion gases exit the combustors and enter a turbine, where the energy of the pressurized combustion gases causes the turbine to rotate. The rotational energy of the turbine is then transmitted, via a shaft, to the compressor and to a generator, for the purpose of generating electricity.
A combustor is typically comprised of at least a pressurized case, a combustion liner, and a transition piece. The combustion liner and transition piece, which contain the high temperature reaction of fuel and air, are subject to thermal degradation. As such, they must be actively cooled to prevent or reduce the degradation rate. In order to actively cool the combustion liner and transition piece, a portion of the compressed air flow is directed through the pressurized case and towards the outer surface of the combustion liner and transition piece, in a generally perpendicular direction, in order to cool these components.
In prior art configurations of gas turbine combustors, exhausted cooling air from the transition piece flows parallel to the surface of the combustion liner mixing with the air being directed through cooling apertures (and towards the outer surface of the combustor liner). Due to the difference in direction of the two air streams, the mixing of the two streams takes place near the surface of the combustor liner. This mixing effect causes the velocity of the air flow perpendicular to surface of the combustor liner (through the cooling apertures) to be reduced. This lowered air flow velocity perpendicular to the surface of the combustor liner leads to less effective cooling of the combustor liner, further accelerating thermal degradation of the combustor liner. Thermal degradation of the liner can lead to premature repair or complete replacement of the liner.
Referring to FIG. 1, a cross sectional perspective view of a prior art gas turbine combustor is shown having a combustion liner 100 encompassed by a flow sleeve 102, forming a flow annulus 104 therebetween. The flow sleeve 102 is provided with a plurality of impingement holes 106, for the purposes of cooling combustion liner 100 on its surface. FIG. 1 also depicts a portion of a gas turbine combustor transition piece 108, which includes an outer mounting flange 110 for coupling the transition piece 108 to the flow sleeve 102 and an inner mounting interface for coupling the transition piece 108 to the combustion liner 100.
Referring now to FIG. 2, a cross sectional view of a portion of the liner 100 and flow sleeve 102 of FIG. 1 is depicted. As discussed above, a generally cylindrical combustion liner 100 and flow sleeve 102 are provided, forming a flow annulus 104 therebetween. Located along the length of flow sleeve 102 is a plurality of impingement holes 106. In a gas turbine combustor, impingement holes 106 are located along a portion of the flow sleeve 102 for providing an impingement flow 112 onto the outer surface of combustion liner 100. Additionally, prior art gas turbine combustors are known to have a cross flow 114 exiting from the transition piece 108 flow annulus and travelling parallel to the outer surface of combustor liner 100. Because the impingement flow 112 and cross flow 114 are generally perpendicular to one another, a substantial portion of cooling impingement flow 112 is turned by the cross flow 114 and is inhibited from reaching the outer surface of the combustor liner 100, as the cross flow 114 significantly reduces the perpendicular velocity component of impingement flow 112.